This invention relates to turbomachinery, and the challenges involved in the production of power or propulsive thrust. In a piece of turbomachinery, like a gas turbine engine, air is pressurized in a compressor then mixed with fuel and burned in a combustor to generate hot combustion gases. The hot combustion gases are expanded within the turbine section where energy is extracted to power the compressor and to produce useful work, such as powering a propeller for an aircraft in flight or turning a generator to produce electricity. The hot core gas travels through a series of turbine stages. A turbine stage may include a row of stationary vanes followed by a row of rotating blades. Each row within the turbine may include a number of airfoils. Each airfoil may be solid, or may contain one or more internal cavities surrounded by an external wall to reduce weight and/or to facilitate a means for internal cooling. The pressure and suction sides of the external wall extend between the leading and trailing edges of the airfoil. Stationary vane airfoils extend span-wise between inner and outer end walls, and the rotating blade airfoils extend span-wise between a platform (typically attached to a rotor disk or other rotary base) and the blade tip.
The stationary vanes and rotating blades of the gas turbine are exposed to high temperature gases, consisting of a mixture of combustion products and cooling and leakage air. In actual practice, the temperature of the gases exiting the combustor will typically exceed the capabilities of the materials used to fabricate the turbine components, necessitating the need for cooling methods to maintain metal temperatures within material specification limits. Generally, the hot gas flows around the pressure and suction side surfaces of the airfoil, from the leading edge to the trailing edge.
However, in the case of a rotating blade, some of the hot gas flows from the pressure side of the airfoil to the suction side through the gap necessarily formed by the rotating blade tip and the adjoining stationary blade outer air seal. Thus, the tip surfaces are exposed to the high temperature environment of the hot gas. During operation of the gas turbine engine, heating of the blade tips by the hot gas may lead to premature thermal distress or failure of the component. It is known in the art that the edges formed by the intersection of the airfoil and tip surfaces are subject to very high heat loads and are therefore predisposed to thermal distress. Because of the complexity and relative high cost of replacing or repairing the blades, it is desirable to prolong the life of the blade tips and respective blades as long as possible.
Air, typically bled off of a compressor stage at a lower temperature and higher pressure than the hot gas passing through the turbine section, may be used to cool the airfoils. The supplied air is generally at higher pressure and lower temperature than that of the hot gases surrounding the blade. Thus, the air extracted from the compressor provides the low temperature sink required for convection heat transfer and film cooling, while the difference in pressure provides the energy required to pass the cooling air through the stationary vane or rotating blade airfoil out to the surrounding gas flow. This use of cooling air permits increased turbine power output by allowing operation of the engine at higher gas temperatures. However, injection of cooling air into the turbine can also reduce gas turbine efficiency.
The gas turbine engine efficiency is, at least in part, dependant upon the extent to which the high temperature gases leak across the gap between the turbine blade tips and the seals or shrouds which surround them. The leakage quantity is typically minimized by positioning the radially-outward blade tip in close proximity to the outer air seal. However, differential thermal elongation and dynamic forces between the blade tip and outer air seal can cause rubbing therebetween. To accommodate this rubbing, abrasive tip treatments and/or squealer ribs, consisting of raised rails extending from the tip cap, may be employed. The squealer ribs are typically exposed to the hot combustion gases on multiple sides and are therefore difficult to cool.
It is therefore desirable to establish better cooling mechanisms along the surfaces of turbine blades, especially near the tip of rotating airfoils. Film cooling is one means to achieve this. A film of cooling air traveling along the surface of the airfoil transfers thermal energy away from the airfoil, increases the uniformity of cooling, and insulates the airfoil from the passing hot gas. However, film cooling is difficult to establish and maintain in the turbulent environment of the gas turbine engine. The presence of non-uniform pressure around the periphery of the airfoil further complicates the film cooling system. The film cooling system must function while discharging coolant to both the high and low pressure sides of the airfoil. In most cases, film cooling air is bled out of apertures extending through the external wall of the airfoil. The term “bled” reflects the small difference in pressure motivating the cooling air out of the internal cavity of the airfoil. However, many known film cooling systems are not efficient, and do not properly address the issues surrounding the cooling of the blade tip, at the edge or interface between the airfoil surfaces and the blade tip cap.
It is desirable therefore, to provide a system and method for cooling a turbomachine blade, and, in particular, to cool the tip region of a rotating airfoil.